1. Field of the Invention
The present invention relates to gas turbine engines and more particularly to stator vanes for use in engines having high turbine inlet temperatures.
2. Description of the Prior Art
The design and construction of gas turbine engines has always required precise engineering effort to ensure the structural integrity of the individual components. One particularly critical area for concern is the turbine nozzle which is formed of a plurality of vanes disposed across the flow path for the high temperature gases in the turbine. During operation of the engine, the flowing gases are redirected by the nozzle onto the rotor blades of a turbine wheel. The temperature of the gases at the inlet to the turbine normally exceeds the allowable temperature limit of the material from which the vanes are fabricated. Consequently, the vanes are cooled to prolong their service life by reducing the metal temperature of the vanes during operation.
Cooling air to the vanes is supplied by the compressor section of the engine. The air is flowed through various conduit means both inwardly and outwardly of the working medium gas path to the turbine section of the engine. A hollow cavity within the airfoil section of each vane receives the cooling air. Air entry ports at both ends of the hollow cavity are in communication with the conduit means. A typical vane utilized in cooled turbines is shown in U.S. Pat. application Ser. No. 531,632 entitled, "Cooled Turbine Vanes" by Leogrande et al, of common assignee herewith. In Leogrande et al an insert is disposed within a hollow cavity at the leading edge of a vane airfoil section. The insert is positioned to direct adequate quantities of cooling air to the leading edge of the airfoil section for film cooling.
Film cooling requires a precise but relatively low pressure differential across flow emitting holes. If the pressure drop is too high, the emitted flow penetrates the passing medium and is deflected downstream with the combustion gases without establishing a film layer on the airfoil surface. On the other hand, if the pressure drop is too small, the hot combustion gases penetrate the cooling air layer to cause destructive heating of the vane material. Because the pressure differential between the cooling air within the vane cavity and the working medium gases at the vane leading edge is relatively small, the amount of flow through each hole is highly sensitive to local pressure deviations within the cavity.
To implement uniform film cooling at the leading edge of the airfoil section, the local pressure deviations in the hollow cavity must be reduced or eliminated. Continuing efforts toward that end are being made.